This invention relates generally to gas turbine engines and more particularly to an aft fan for a gas turbine engine.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. In a turbofan engine, which typically includes a fan placed at the front of the core engine, a high pressure turbine powers the compressor of the core engine. A low pressure turbine is disposed downstream from the high pressure turbine for powering the fan.
Some prior art engine configurations incorporate an aft fan stage integral with a turbine rotor. There are several potential advantages for this “fan on turbine” configuration, which eliminates the drive shaft required in a front-fan engine. It is desired to have such a fan-on-turbine with a design pressure ratio (i.e. the ratio of total pressure at the fan exit to the total pressure at the fan inlet) of about 2.5 or greater. Unfortunately, the high tip speed required for a single fan stage to produce this pressure ratio is contrary the AN2 and radius ratio constraints dictated by accepted turbine design practice.
Accordingly, there is a need for a fan-on-turbine configuration which achieves a high pressure ratio.